The disclosure relates to gas turbine engines. More particularly, the disclosure relates to thermal barrier coating systems.
Gas turbine engine combustor and turbine components are exposed to extreme heat during engine operation. They are also exposed to the mechanical action of the gas flow and to chemical reaction from combustion gas and entrained matter. Exemplary such hot section components include combustor panels, vanes, blades, air seals, and the like. Accordingly, it is customary to coat such components with thermal barrier coating (TBC) systems.
Exemplary thermal barrier coating systems include two-layer thermal barrier coating systems. An exemplary system includes a MCrAlY bond coat (where M identifies one or more of Fe, Ni, and Co) and a yttria-stabilized zirconia (YSZ) thermal barrier coat (TBC). While barrier coat layer is being deposited or during an initial heating cycle, a thermally grown oxide (TGO) layer (e.g., alumina) forms atop the bond coat layer. As time-at-temperature and the number of cycles increase, this TGO interface layer grows in thickness. U.S. Pat. Nos. 4,405,659 and 6,060,177 disclose exemplary systems.
Especially when used in sandy desert environments, engine components are subject to a particular form of fouling/damage known as molten sand attack or CMAS. CMAS is an abbreviation for “calcium-magnesium-aluminum-silicon (silicate)”. Specific CMAS oxides include CAO, MGO, Al2O3, and SiO2. CMAS components may form a eutectic with a relatively low melting point (e.g., approximately 1240 C). The molten CMAS material infiltrates into porous coatings (e.g., between the columns of columnar ceramic). This can alter the chemical composition of the coating and/or cause structural failure of the coating. Efforts to address CMAS have centered on improved barrier coatings. For example, U.S. Pat. Nos. 5,660,885, 5,871,820, 5,914,189, 6,720,038, 6,627,323, 6,465,090, all reference coatings relative to CMAS.